Heat shield with axial retention lock

ABSTRACT

A heat shield assembly for an engine case of a gas turbine engine may include a heat shield and a support lock. The heat shield may have an annular shape. The heat shield may define an aperture extending through the heat shield. The support lock may have a tab extending radially outward from a distal surface of the support lock. The aperture in the heat shield may be configured to retain the tab of the support lock.

FIELD

The present disclosure relates to gas turbine engines, and, morespecifically, to a combustor section and engine case.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. In general, duringoperation, air is pressurized in the compressor section and is mixedwith fuel and burned in the combustor section to generate hot combustiongases. The hot combustion gases flow through the turbine section, whichextracts energy from the hot combustion gases to power the compressorsection and other gas turbine engine loads. The compressor sectiontypically includes low pressure and high pressure compressors, and theturbine section includes low pressure and high pressure turbines.

The combustor is typically coupled to an engine case of the gas turbineengine. The engine case may include a diffuser case, which circumscribesthe compressor section. The diffuser case and fittings may be subjectedto relatively high temperatures due to heat convectively transferredfrom the combustor to the diffuser case. Thermal loads in the diffusercase may cause thermal gradients that may stress, deform, fracture,and/or degrade portions of the diffuser case over time. A flange of thediffuser case may experience thermal gradients of 500° F. (260° C.) to600° F. (315° C.). The thermal gradients cause stress that may shortenthe operational life of engine case components. During operation, thethermal load on an engine case may increase the overall length of theengine case. This thermal growth may contribute to misalignment ofengine components.

SUMMARY

A heat shield assembly for an engine case of a gas turbine engine isdescribed herein, in accordance with various embodiments. The heatshield assembly may include a heat shield and a support lock. The heatshield may have an annular shape. The heat shield may define an apertureextending through the heat shield. The support lock may have a tabextending radially outward from a distal surface of the support lock.The aperture in the heat shield may be configured to retain the tab ofthe support lock.

In various embodiments, a distal surface of the support lock may beconfigured to be coupled to the inner surface of the heat shield. Theaperture may be disposed in proximity to an aft end of the heat shield.The tab may be disposed at a forward end of the support lock. The outersurface of the heat shield may form a seal with the engine case. Theouter surface of the heat shield and an inner surface of the engine casemay define a gap. In various embodiments, the engine case may be adiffuser case.

A combustor section of a gas turbine engine is also provided. Thecombustor section may include a combustor and a diffuser case disposedabout the combustor. A heat shield may have an annular shape and may bedisposed between the combustor and the diffuser case. The heat shieldmay define an aperture extending through the heat shield.

In various embodiments, the heat shield may be configured to couple tothe diffuser case. The heat shield may be disposed circumferentiallyalong an inner surface of the diffuser case. An outer surface of theheat shield may form a seal with the inner surface of the diffuser case.The outer surface of the heat shield and the inner surface of thediffuser case may define a gap. A support lock may have a tab extendingradially outward from a distal surface of the support lock. The aperturein the heat shield may be configured to retain the tab of the supportlock. The tab may be disposed at a forward end of the support lock. Thetab and the aperture may be configured to retain the heat shield in anaxial direction.

A gas turbine engine is also provided. The gas turbine engine mayinclude a combustor and a diffuser case disposed about the combustor. Aheat shield assembly may be disposed between the combustor and diffusercase. The heat shield assembly may include a heat shield having anannular shape. The heat shield may define an aperture extending throughthe heat shield. The heat shield assembly may further include a supportlock having a tab extending radially outward from a distal surface ofthe support lock. The aperture in the heat shield may be configured toretain the tab of the support lock.

In various embodiments, the heat shield may be disposedcircumferentially along an inner surface of the diffuser case. An outersurface of the heat shield may form a seal with the inner surface of thediffuser case. The outer surface of the heat shield and the innersurface of the diffuser case may define a gap.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated otherwise.These features and elements as well as the operation thereof will becomemore apparent in light of the following description and the accompanyingdrawings. It should be understood, however, the following descriptionand drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter of the present disclosure is particularly pointed outand distinctly claimed in the concluding portion of the specification. Amore complete understanding of the present disclosure, however, may bestbe obtained by referring to the detailed description and claims whenconsidered in connection with the figures, wherein like numerals denotelike elements.

FIG. 1 illustrates a cross-sectional view of an exemplary gas turbineengine, in accordance with various embodiments;

FIG. 2 illustrates a cross-sectional view of a combustor and a case of agas turbine engine including a heat shield and an axial retentionfeature, in accordance with various embodiments;

FIG. 3A illustrates a cross-sectional view of a diffuser case flangehaving a heat shield, in accordance with various embodiments;

FIG. 3B illustrates a perspective view of a heat shield, in accordancewith various embodiments;

FIG. 3C illustrates a perspective view of a support lock having an axialretention feature, in accordance with various embodiments; and

FIG. 3D illustrates a perspective view of a heat shield and supportlock, in accordance with various embodiments.

DETAILED DESCRIPTION

All ranges and ratio limits disclosed herein may be combined. It is tobe understood that unless specifically stated otherwise, references to“a,” “an,” and/or “the” may include one or more than one and thatreference to an item in the singular may also include the item in theplural.

The detailed description of various embodiments herein makes referenceto the accompanying drawings, which show various embodiments by way ofillustration. While these various embodiments are described insufficient detail to enable those skilled in the art to practice thedisclosure, it should be understood that other embodiments may berealized and that logical, chemical, and mechanical changes may be madewithout departing from the spirit and scope of the disclosure. Thus, thedetailed description herein is presented for purposes of illustrationonly and not of limitation. For example, the steps recited in any of themethod or process descriptions may be executed in any order and are notnecessarily limited to the order presented. Furthermore, any referenceto singular includes plural embodiments, and any reference to more thanone component or step may include a singular embodiment or step. Also,any reference to attached, fixed, connected, or the like may includepermanent, removable, temporary, partial, full, and/or any otherpossible attachment option. Additionally, any reference to withoutcontact (or similar phrases) may also include reduced contact or minimalcontact. Surface shading lines may be used throughout the figures todenote different parts but not necessarily to denote the same ordifferent materials.

As used herein, “aft” refers to the direction associated with the tail(e.g., the back end) of an aircraft, or generally, to the direction ofexhaust of the gas turbine engine. As used herein, “forward” refers tothe direction associated with the nose (e.g., the front end) of anaircraft, or generally, to the direction of flight or motion.

As used herein, “distal” refers to the direction radially outward, orgenerally, away from the axis of rotation of a turbine engine. As usedherein, “proximal” refers to a direction radially inward, or generally,towards the axis of rotation of a turbine engine.

In various embodiments and with reference to FIG. 1, a gas turbineengine 20 is provided. Gas turbine engine 20 may be a two-spool turbofanthat generally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mayinclude, for example, an augmentor section among other systems orfeatures. In operation, fan section 22 can drive coolant (e.g., air)along a bypass flow-path B while compressor section 24 can drive coolantalong a core flow-path C for compression and communication intocombustor section 26 then expansion through turbine section 28. Althoughdepicted as a turbofan gas turbine engine 20 herein, it should beunderstood that the concepts described herein are not limited to usewith turbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

Gas turbine engine 20 may generally comprise a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A-A′ relative to an engine static structure 36 orengine case via several bearing systems 38, 38-1, and 38-2. Enginecentral longitudinal axis A-A′ is oriented in the z direction on theprovided xyz axis. It should be understood that various bearing systems38 at various locations may alternatively or additionally be provided,including for example, bearing system 38, bearing system 38-1, andbearing system 38-2.

Low speed spool 30 may generally comprise an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. Inner shaft 40 may be connected to fan 42 through a gearedarchitecture 48 that can drive fan 42 at a lower speed than low speedspool 30. Geared architecture 48 may comprise a gear assembly 60enclosed within a gear housing 62. Gear assembly 60 couples inner shaft40 to a rotating fan structure. High speed spool 32 may comprise anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 may be located between high pressurecompressor 52 and high pressure turbine 54. An outer diffuser case 70 ofengine static structure 36 may enclose the combustor 56. A high pressureturbine (HPT) case 72 of engine static structure 36 may enclose highpressure turbine 54. An aft end of outer diffuser case 70 may beattached to a forward end of HPT case 72 at an attachment interface 74.A mid-turbine frame 57 of engine static structure 36 may be locatedgenerally between high pressure turbine 54 and low pressure turbine 46.Mid-turbine frame 57 may support one or more bearing systems 38 inturbine section 28. Inner shaft 40 and outer shaft 50 may be concentricand rotate via bearing systems 38 about the engine central longitudinalaxis A-A′, which is collinear with their longitudinal axes. As usedherein, a “high pressure” compressor or turbine experiences a higherpressure than a corresponding “low pressure” compressor or turbine.

The core airflow C may be compressed by low pressure compressor 44 thenhigh pressure compressor 52, mixed and burned with fuel in combustor 56,then expanded over high pressure turbine 54 and low pressure turbine 46.Turbines 46, 54 rotationally drive the respective low speed spool 30 andhigh speed spool 32 in response to the expansion.

Gas turbine engine 20 may be, for example, a high-bypass ratio gearedaircraft engine. In various embodiments, the bypass ratio of gas turbineengine 20 may be greater than about six (6). In various embodiments, thebypass ratio of gas turbine engine 20 may be greater than ten (10). Invarious embodiments, geared architecture 48 may be an epicyclic geartrain, such as a star gear system (sun gear in meshing engagement with aplurality of star gears supported by a carrier and in meshing engagementwith a ring gear) or other gear system. Geared architecture 48 may havea gear reduction ratio of greater than about 2.3 and low pressureturbine 46 may have a pressure ratio that is greater than about five(5). In various embodiments, the bypass ratio of gas turbine engine 20is greater than about ten (10:1). In various embodiments, the diameterof fan 42 may be significantly larger than that of the low pressurecompressor 44, and the low pressure turbine 46 may have a pressure ratiothat is greater than about five (5:1). Low pressure turbine 46 pressureratio may be measured prior to inlet of low pressure turbine 46 asrelated to the pressure at the outlet of low pressure turbine 46 priorto an exhaust nozzle. It should be understood, however, that the aboveparameters are exemplary of various embodiments of a suitable gearedarchitecture engine and that the present disclosure contemplates othergas turbine engines including direct drive turbofans.

With reference to FIG. 2, a combustor section 26 is shown, in accordancewith various embodiments. Combustor section 26 generally includescombustor 56, which may be coupled to outer diffuser case 70 by adiffuser mount assembly 100. Combustor 56 generally includes acombustion chamber 102 defined by a combustor wall 104. Combustor 56 maybe encased by outer diffuser case 70 having an annular geometry anddisposed about combustor 56. Combustor 56 may be further encased by aninner diffuser case 106. Inner diffuser case 106 is spaced radiallyinward from combustor wall 104 to define an inner plenum 108. Outerdiffuser case 70 is spaced radially outward from combustor wall 104 todefine an outer plenum 110.

Referring briefly to FIG. 1, combustor 56 may be disposed downstream ofthe compressor section 24 to receive compressed airflow therefrom.Referring now to FIG. 2, Combustion chamber 102 contains the combustionproducts that flow axially toward turbine section 28. Gas leaving highpressure compressor 52 may flow into combustion chamber 102 to supplycombustor 56 with air for combustion. Uncombusted gas may be mixed withfuel and burned in combustion chamber 102. Combusted gas in combustor 56may reach or exceed temperatures of up to 3,500° F. (1,925° C.) orhigher. Heat may radiate from combustor 56 to other nearby componentswhich may cause the nearby components to increase in temperature.

With momentary reference to FIG. 1, turbine section 28 receivescombusted gas or exhaust from combustor section 26. In variousembodiments, turbine section 28 may include multiple rows of vanes andmultiple rows of blades that can rotate about an axis with respect tothe vanes. Combusted gas from the combustor section 26 is channeled tothe turbine section where it can be directed through the turbine vanesand blades. With reference to FIGS. 1 and 2, high pressure turbine 54may include a plurality of vanes, such as vane 112, and a plurality ofblades. Vane 112 may operate as a first stage high pressure turbine vaneof high pressure turbine 54. Vane 112 and combustor 56 may attach toouter diffuser case 70 by a combustor vane support 114. Combustor 56 maybe further secured to outer diffuser case 70 and combustor vane support114 by a plurality of support locks 116. In various embodiments,diffuser mount assembly 100 includes a plurality of support locks 116disposed at regular intervals circumferentially around combustor 56.Each support lock 116 and combustor vane support 114 may be configuredto receive a fastener 118. A fastener 118 may be passed throughcombustor vane support 114 and support lock 116 to mechanically couplesupport lock 116 to combustor vane support 114.

In various embodiments, outer diffuser case 70 may be attached to HPTcase 72. Outer diffuser case 70 may include a diffuser case flange 120extending radially from outer diffuser case 70 at or near the aft end ofouter diffuser case 70. HPT case 72 may include an HPT case flange 122extending radially from HPT case 72 at or near the forward end of HPTcase 72. Diffuser case flange 120 and HPT case flange 122 may matetogether at attachment interface 74. Flanges 120, 122 may be in directlyabutting engagement with each other at attachment interface 74 and maybe secured by a plurality of fasteners 124. Fasteners 124 may be passedthrough flanges 120, 122 to mechanically couple outer diffuser case 70and HPT case 72. Fasteners 124 may include rivets, bolts, or othersuitable fasteners to couple outer diffuser case 70 and HPT case 72along flanges 120, 122.

Heat radiating from combustor 56 may introduce thermal loads on outerdiffuser case 70 and diffuser case flange 120. For example, heat mayconvectively transfer from combustor 56 to outer diffuser case 70 and todiffuser case flange 120, resulting in a thermal gradient in diffusercase flange 120. A heat shield 126 may be configured to block heatradiating from combustor 56 from directly impinging on outer diffusercase 70 and on diffuser case flange 120. Heat shield 126 may be disposedbetween combustor 56 and outer diffuser case 70. In various embodiments,a heat shield 126 may extend circumferentially along an inner wall ofouter diffuser case 70. Heat shield 126 may help reduce the thermalgradients in diffuser mount flange 120, thereby reducing stress ondiffuser case flange 120.

With reference to FIG. 3A, a cross-sectional view of diffuser mountassembly 100 is shown with heat shield 126, in accordance with variousembodiments. Heat shield 126 may be used to insulate outer diffuser case70 and diffuser case flange 120 from convective heat transfer fromcombustor 56 through gas flowing through outer plenum 110. In variousembodiments, a forward end 130 of heat shield 126 may contact outerdiffuser case 70 at a contact surface 132. An aft end 134 of heat shield126 may further contact outer diffuser case 70 at a contact surface 136.Heat shield 126 may form a seal with outer diffuser case 70 at contactsurface 132 and contact surface 136. Contact surfaces 132, 136 may eachform an annular seal disposed circumferentially along an inner surface140 of outer diffuser case 70. A portion of heat shield 126 may beconfigured to be separated from outer diffuser case 70 by a gap 144. Invarious embodiments, an outer surface 142 of heat shield 126 and aninner surface 140 of outer diffuser case 70 define gap 144, which isdisposed between heat shield 126 and outer diffuser case 70. Gap 144 mayextend axially and circumferentially along heat shield 126 betweencontact surfaces 132, 136. Gap 144 may be configured such that aconductive thermal path does not exist between heat shield 126 and outerdiffuser case 70 at gap 144. Gap 144 may be configured to minimizeconvective heat transfer between heat shield 126 and outer diffuser case70. Accordingly, heat shield 126 may be configured to minimizeconvective heat transfer between combustor 56 and outer diffuser case70, thereby decreasing the temperature of diffuser case flange 120. Thereduced temperature of diffuser case flange 120 reduces thermalgradients and stress experienced by diffuser case flange 120 andincreases the operational life of diffuser case flange 120.

Core airflow C generally flows through combustor section 26 in thedirection of arrows 146. During a surge event, core airflow C maybackflow or may travel in a direction other than in the direction ofarrows 146. Backflow of core airflow C can disrupt the position ofengine components and may lead to liberation of heat shield 126 (e.g.,heat shield 126 may loosen or detach from outer diffuser case 70),increasing the risk of damage to the surrounding engine structure.Thermal growth of engine components, such as the engine case, may alsolead to liberation of heat shield 126. An axial retention feature 150may be included to prevent axial liberation of heat shield 126, due tothermal growth and/or during a surge event. In various embodiments,axial retention feature 150 may include a plurality of apertures 152formed in heat shield 126. Apertures 152 may be configured as openingsformed completely through heat shield 126, wherein apertures 152 mayextend from outer surface 142 to an inner surface 154 of heat shield126. Inner sidewalls of heat shield 126 may define apertures 152.

Axial retention feature 150 may further include a tab 156 of supportlock 116. In various embodiments, each of the plurality of apertures 152may be configured to receive a tab 156 of support lock 116. Tab 156disposed within aperture 152 may form a tight fit such that gap 144remains sealed between outer surface 142 of heat shield 126 and innersurface 140 of outer diffuser case 70. Tab 156 may extend completelythough aperture 152 and may further extend radially outward beyond outersurface 142 of heat shield 126 and into gap 144. In various embodiments,tab 156 extending into gap 144 may not extend completely through gap 144and may not contact inner surface 140 of outer diffuser case 70. Tab 156may remain spaced apart from outer diffuser case 70 such that a thermalconduction path is not formed between tab 156 and outer diffuser case70. Tab 156 remaining spaced apart from outer diffuser case 70 improvesthe thermal shielding of outer diffuser case 70 and diffuser case flange120.

In various embodiments, tab 156 of support lock 116 may fit throughaperture 152 of heat shield 126 to form axial retention feature 150.Support lock 116 may be secured to combustor vane support 114 byfastener 118, and tab 156 of support lock 116 secures heat shield 126 tocombustor vane support 114 with respect to axial motion. Further,combustor vane support 114 may be coupled to outer diffuser case 70.Accordingly, tab 156 may be coupled to outer diffuser case 70 throughsupport lock 116, fastener 118, and combustor vane support 114. Theconfiguration of tab 156 coupled to outer diffuser case 70 and toaperture 152 of heat shield 126 may prevent heat shield 126 from movingaxially in the z direction, for example, with respect to outer diffusercase 70. Thus, axial retention feature 150 may retain heat shield 126 inthe axial direction or reduce movement of heat shield 126 in the axialdirection, or z direction. Tab 156 extending completely through heatshield 126 may provide improved axial retention during thermal growth ora surge event.

With reference to FIG. 3B, a perspective view of heat shield 126 isshown in accordance with various embodiments. In various embodiments,heat shield 126 may be circumferentially continuous (e.g., annular orring shaped). Heat shield 126 having annular geometry may interface withand fit within outer diffuser case 70, which may also have annulargeometry. Heat shield 126 includes an outer surface 142 and an innersurface 154. In various embodiments, heat shield 126 may include aplurality of apertures 152 disposed at regular intervals around thecircumference of heat shield 126. Apertures 152 may be configured toalign with support locks 116. A number of apertures 152 may correspondto the number of support locks 116. Apertures 152 may be formed inproximity to aft end 134 of heat shield 126.

In various embodiments, heat shield 126 may be manufactured using sheetmetal, forging, casting, additive manufacturing, machining or the like.Apertures 152 may further be formed by milling, electrochemicalmachining (ECM), or electrostatic discharge machining (EDM) as desired,for example. It is desirable for the material of heat shield 126 to beresistant to heat. In that regard and in various embodiments, heatshield 126 may include a high performance nickel-based super alloy orother suitable material.

With reference to FIG. 3C, a perspective view of support lock 116 isshown with an axial retention feature, in accordance with variousembodiments. In various embodiments, support lock 116 may be configuredto interface with combustor vane support 114 at a mating surface 160.Support lock 116 may be configured to receive a fastener, for example,within opening 162. Support lock 116 may be further configured with adistal surface 164 having tab 156 extending radially outward from distalsurface 164. Tab 156 is depicted with a rectangular shape having a sidesurface 166 and a top surface 168. Side surface 166 may extend radiallyoutward from distal surface 164. A forward side surface of tab 156 maybe coplanar or continuous with a forward surface 170 of support lock116. Tab 156 may operate as a first axial retention feature of supportlock 116 and may be configured to complement and interface with a secondaxial retention feature, such as aperture 152, of heat shield 126. Whilethe first axial retention feature, or tab 156, of support lock 116 isdepicted as rectangular in shape, it is to be understood that any shapecomplementary to the second axial retention feature, or aperture 152, ofheat shield 126 may be used.

In various embodiments, support lock 116 may be manufactured by forging,casting, additive manufacturing, machining such as ECM, EDM, or thelike. In various embodiments, support lock 116 may include an austeniticnickel-chromium-based alloy, a high performance nickel-based superalloy, or other suitable material.

With reference to FIG. 3D, a perspective view of a heat shield assembly172 is shown with axial retention feature 150, in accordance withvarious embodiments. In various embodiments, heat shield assembly 172may include heat shield 126 and support lock 116. In variousembodiments, heat shield assembly 172 may further include a plurality ofaxial retention features 150. Each axial retention feature 150 maycomprise an aperture 152 and a tab 156. In various embodiments, tab 156may extend radially outward and integrally from distal surface 164 ofsupport lock 116. Tab 156 may engage with aperture 152 to prevent axialliberation of heat shield 126. Tab 156 of support lock 116 may fit intoaperture 152 of heat shield 126 such that tab 156 extends completelythrough a thickness of heat shield 126. An interference fit may beformed with tab 156 disposed within aperture 152 to help retain heatshield 126 in a position along the z axis. In various embodiments, heatshield 126 may be integral to (e.g., forged or manufactured as part of)support lock 116. Apertures 152 and tab 156 may be disposed in proximityto aft end 134 of heat shield 126. Tab 156 may be disposed at a forwardend of support lock 116. A position of tab 156 and apertures 152 isconfigured to provide both axial retention and thermal shielding. Aconfiguration with tab 156 extending through heat shield 126 may improveaxial retention. A configuration with tab 156 disposed at a forward endof support lock 116 may improve thermal shielding by reducing heattransfer from support lock 116 and tab 156 to outer diffuser case 70.

Benefits and other advantages have been described herein with regard tospecific embodiments. Furthermore, the connecting lines shown in thevarious figures contained herein are intended to represent exemplaryfunctional relationships and/or physical couplings between the variouselements. It should be noted that many alternative or additionalfunctional relationships or physical connections may be present in apractical system. However, the benefits, advantages, and any elementsthat may cause any benefit or advantage to occur or become morepronounced are not to be construed as critical, required, or essentialfeatures or elements of the disclosure. The scope of the disclosure isaccordingly to be limited by nothing other than the appended claims, inwhich reference to an element in the singular is not intended to mean“one and only one” unless explicitly so stated, but rather “one ormore.” Moreover, where a phrase similar to “at least one of A, B, or C”is used in the claims, it is intended that the phrase be interpreted tomean that A alone may be present in an embodiment, B alone may bepresent in an embodiment, C alone may be present in an embodiment, orthat any combination of the elements A, B and C may be present in asingle embodiment; for example, A and B, A and C, B and C, or A and Band C.

Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “various embodiments”, “oneembodiment”, “an embodiment”, “an example embodiment”, etc., indicatethat the embodiment described may include a particular feature,structure, or characteristic, but every embodiment may not necessarilyinclude the particular feature, structure, or characteristic. Moreover,such phrases are not necessarily referring to the same embodiment.Further, when a particular feature, structure, or characteristic isdescribed in connection with an embodiment, it is submitted that it iswithin the knowledge of one skilled in the art to affect such feature,structure, or characteristic in connection with other embodimentswhether or not explicitly described. After reading the description, itwill be apparent to one skilled in the relevant art(s) how to implementthe disclosure in alternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element herein is to be construed under theprovisions of 35 U.S.C. 112(f), unless the element is expressly recitedusing the phrase “means for.” As used herein, the terms “comprises”,“comprising”, or any other variation thereof, are intended to cover anon-exclusive inclusion, such that a process, method, article, orapparatus that comprises a list of elements does not include only thoseelements but may include other elements not expressly listed or inherentto such process, method, article, or apparatus.

What is claimed is:
 1. A heat shield assembly for an engine case of agas turbine engine, comprising: a heat shield having an annular shape,wherein the heat shield defines an aperture extending through the heatshield; and a support lock having a tab extending radially outward froma distal surface of the support lock, wherein the aperture in the heatshield is configured to retain the tab of the support lock.
 2. The heatshield assembly of claim 1, wherein the distal surface of the supportlock is configured to be coupled to an inner surface of the heat shield.3. The heat shield assembly of claim 2, wherein the aperture is disposedin proximity to an aft end of the heat shield.
 4. The heat shieldassembly of claim 3, wherein the tab is disposed at a forward end of thesupport lock.
 5. The heat shield assembly of claim 4, wherein an outersurface of the heat shield forms a seal with the engine case.
 6. Theheat shield assembly of claim 5, wherein the outer surface of the heatshield and an inner surface of the engine case define a gap.
 7. The heatshield assembly of claim 1, wherein the engine case comprises a diffusercase.
 8. A combustor section of a gas turbine engine, comprising: acombustor; a diffuser case disposed about the combustor; and a heatshield having an annular shape and disposed between the combustor andthe diffuser case, wherein the heat shield defines an aperture extendingthrough the heat shield.
 9. The combustor section of claim 8, whereinthe heat shield is configured to couple to the diffuser case.
 10. Thecombustor section of claim 8, wherein the heat shield is disposedcircumferentially along an inner surface of the diffuser case.
 11. Thecombustor section of claim 10, wherein an outer surface of the heatshield forms a seal with the inner surface of the diffuser case.
 12. Thecombustor section of claim 11, wherein the outer surface of the heatshield and the inner surface of the diffuser case define a gap.
 13. Thecombustor section of claim 8, further comprising a support lock having atab extending radially outward from a distal surface of the supportlock.
 14. The combustor section of claim 13, wherein the aperture in theheat shield is configured to retain the tab of the support lock.
 15. Thecombustor section of claim 14, wherein the tab is disposed at a forwardend of the support lock.
 16. The combustor section of claim 15, whereinthe tab and the aperture are configured to retain the heat shield in anaxial direction.
 17. A gas turbine engine, comprising: a combustor; adiffuser case disposed about the combustor; and a heat shield assemblydisposed between the combustor and the diffuser case, the heat shieldassembly comprising: a heat shield having an annular shape, wherein theheat shield defines an aperture extending through the heat shield, and asupport lock having a tab extending radially outward from a distalsurface of the support lock, wherein the aperture in the heat shield isconfigured to retain the tab of the support lock.
 18. The gas turbineengine of claim 17, wherein the heat shield is disposedcircumferentially along an inner surface of the diffuser case.
 19. Thegas turbine engine of claim 18, wherein an outer surface of the heatshield forms a seal with the inner surface of the diffuser case.
 20. Thegas turbine engine of claim 19, wherein the outer surface of the heatshield and the inner surface of the diffuser case define a gap.